1. Field of the Invention
The present invention relates to a rocket propulsion system configured for intermittent operation and steerability.
2. State of the Art
In a multi-stage rocket motor propulsion system, an upper stage rocket motor carrying a payload typically sits atop one or more lower stage rocket motors. Each stage of a multistage rocket motor-propelled vehicle is separated and discarded once the fuel thereof has been consumed. Discarding successive rocket motor stages reduces the weight of the fuselage and increases the mass ratio of the rocket. This approach enables greater range, greater capability in boosting heavy payloads or a combination of such advantages. The upper stage rocket motor comes into operation at high altitude after the one or more lower stage rocket motors have separated from the vehicle.
Rocket motor propulsion systems for rockets and missiles are generally either liquid propellant or solid propellant-based, although so-called “hybrid” propulsion systems using both solid and liquid propellant components are known.
Liquid propellant rocket engines mix liquid fuel and liquid oxidizer in a combustion chamber in a specific proportion and at flow rates designed to cause the liquid to spontaneously combust. Propulsion thrust occurs as the gaseous combustion products are expelled from the rocket motor's exhaust nozzle. Liquid propellant rocket engines can be controlled, stopped and restarted. Disadvantages of using a liquid propellant include the volatility of the liquid fuel, the high level of care required during storage and handling thereof and the requirement that the propellant be loaded into the rocket motor immediately prior to launch.
Solid propellant rocket motors employ a propellant comprising a solid fuel charge or “grain” which burns to generate exhaust gases and other combustion products, which are expelled through a nozzle of the rocket motor to provide thrust. Once a grain of solid propellant is ignited it is difficult to extinguish and the entire grain is ordinarily consumed after ignition. Additionally, effecting variation of thrust is more difficult in solid propellant than in liquid propellant rocket engines. However, simple structural design of solid propellant rocket motors and ease of storage of the solid propellant are advantages of the solid propellant motor.
One method of fabricating a solid propellant rocket motor having the capability for shut down and reignition is to provide multiple propellant masses or “pulses.” Each “pulse” may be one layer or zone of solid propellant disposed in a combustion chamber, with a flame-inhibiting barrier separating the layers. The flame-inhibiting barrier is made of a material that will confine the propellant burning to a single layer or zone, yet is selectively destructible so that the next adjacent layer may be ignited. The burning of each layer of solid propellant produces a thrust in the form of a discrete pulse. The number of pulses, as well as the burn time of each pulse, commonly termed the “duty cycle,” must be sized prior to fabrication of the rocket motor. These requirements limit operational flexibility of the rocket motor. In addition, thrust may only be terminated once a pulse burns out.
One method for termination of the combustion of the propellant in a solid propellant rocket motor is rapid depressurization of the pressure vessel in which such combustion takes place. Depressurization may be effected by explosive ejection of a plug sealing an opening in the wall of the pressure vessel. An alternate depressurization method is to explosively sever the nozzle assembly from the aft end of the rocket to open up a substantially larger exit port, causing rapid depressurization. The disadvantage of these approaches is the associated extremely high acceleration jolt, whether in a forward, rearward, or axially offset direction. Further, these approaches limit, if not destroy, any subsequent operability of the rocket motor.
Attitude control, in the form of influencing the pitch, yaw, and/or roll of the rocket assembly in flight, may be accomplished with a thrust vector control (TVC) system or a separate attitude control system (ACS).
A TVC system may comprise an axial thrust nozzle rotationally positionable at a desired angle within a range offset from the longitudinal axis of the rocket motor to alter the vector at which the combustion products exit the rocket motor. Repositioning of the nozzle alters the direction of the forces acting on the vehicle in which the rocket motor is installed to alter the vehicle's direction of flight. Single, moveable TVC nozzles provide adequate control over the rocket assembly's yaw and pitch, but do not provide any significant degree of roll control.
Multiple rocket engines or gas generators and associated thrusters are often employed to control attitude. The rocket engines or thrusters are offset from the longitudinal axis of the rocket motor assembly so that firing of selected ones or groups of the engines or powering of selected ones or groups of thrusters enables attitude control over the rocket motor assembly. Use of a separate ACS in combination with one or more axial thrust engines or thrusters increases the weight of the rocket motor assembly due to the additional hardware. A separate ACS may use a solid-propellant gas generator directly connected to a manifold providing a selective hot gas flow to nozzle valve clusters.
Roll control may be provided by the ACS or through the inclusion of a separate roll control system (RCS). Separate gas generators and thrusters may be provided for the RCS.
In view of the above-enumerated deficiencies in the state of the art with respect to both liquid propellant rocket engines and solid propellant rocket motors, it would be desirable to develop a rocket motor which uses a solid propellant which may be selectively extinguished and reignited a plurality of times. A rocket motor offering integral attitude control capabilities, including pitch, yaw and roll control capabilities, would also be desirable.